Method of manufacturing a panel for an aircraft propulsion unit nacelle

ABSTRACT

A method for manufacturing a panel includes disposing one or several supports in a respective cavity(ies) in a honeycomb structure having a first skin and a second skin clasping the honeycomb structure. The supports are made of a fugitive material such as a thermoplastic and are auxetic so that, under the effect of an increase in temperature, their dimension between the first and the second skins remains below a predetermined value.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to and the benefit of FR 18/57692 filedon Aug. 27, 2018. The disclosure of the above application isincorporated herein by reference.

FIELD

The present disclosure relates to the field of the manufacture of panelsfor an aircraft propulsion unit nacelle, in particular panels of thetype comprising a honeycomb structure clasped between two skins.

BACKGROUND

The statements in this section merely provide background informationrelated to the present disclosure and may not constitute prior art.

There are known in the state of the art methods for manufacturing panelsby assembling composite or metal skins with a honeycomb structure—seefor example WO 2017/017369 A1. In this document, the cavities of thehoneycomb structure are filled with a fugitive material intended toprevent the collapse of the skins at the level of the cavities.

Nonetheless, the increase in temperature during sintering tends toexpand the fugitive material, before the removal thereof. This can causedeformation of the skins under the effect of the forces exerted thereonby the expanded fugitive material.

SUMMARY

In one form of the present disclosure a method for manufacturing a panelfor an aircraft propulsion unit nacelle capable of preserving moreeffectively the inner volume of cavities during an increase intemperature, in particular during a sintering step, is provided.

In some variations the panel comprises a honeycomb structure formingcavities as well as a first skin and a second skin clasping thehoneycomb structure. In one variation the cavities may be Helmholtzcavities.

In some variations the method comprises a step of disposing one orseveral support(s) in respective cavities of the honeycomb structure.The supports are made of a fugitive material such as a thermoplastic.

A fugitive material allows removing the supports of the cavities byheating, their removal from cavities being able to be carried out by anyconventional technique or process known to those skilled in the art, forexample via perforations made in one of the skins, or via channelsprovided to this end.

According to the teachings of the present disclosure, the supports areauxetic so that, under the effect of an increase in temperature, theirdimension between the first and the second skins remains below apredetermined value.

In some variations of the present disclosure the predetermined valuewill be selected so that, under the effect of an increase intemperature, the dimension of the supports between the first and thesecond skins remains identical or presents a small or negligibleincrease with regard to the forces exerted by these supports against theskin under the effect of the expansion of these supports.

It is thus possible to avoid or limit any deformation of the skins underthe effect of the thermal expansion of the supports.

In one form, the supports may have a hexagonal re-entrant shape. Besidesthe auxetic property conferred by such a shape, it also allows reducingthe amount of material constituting these supports, and thereforelimiting the amount of material to be removed during the removal of thesupports.

In one form, the method may comprise a step of making the first skin andthe honeycomb structure by additive manufacturing.

In such a form, the step of making the first skin and the honeycombstructure comprises a deposition of a compound comprising a TiAl-basedpowdery intermetallic alloy.

TiAl is a material having good performance at high temperature andhaving a relatively low density.

The step of making the first skin and the honeycomb structure maycomprise consolidation of the compound by sintering.

In some variations any step of melting the compound is not implemented.In other words, the consolidation of the compound can therefore beobtained only by heating this compound below its melting temperature.

The absence of melting of the powder particles reduces or avoids TiAlhot cracking.

More generally, the mechanical properties of the panel are thusimproved.

Despite the poor ductility of TiAl, the deposition of a compoundcomprising this material and its consolidation by sintering allowssimple and inexpensive manufacturing of a panel with a complex geometrywhile conferring on this panel a reduced mass and good performance athigh temperature.

In some forms, the second skin may be made by depositing a compoundcomprising a TiAl-based powdery intermetallic alloy or bepre-manufactured and affixed against the honeycomb structure.

Regarding the supports, these may also be made by additive manufacturingsimultaneously with the making of the first skin and/or of the honeycombstructure.

For this purpose, it is possible, for example, to use one technique orprocess to deposit the compound and another technique or process fordepositing the material in which the supports are made, and in somevariations of the present disclosure these different depositingtechniques or processes are arranged for working in parallel.

Such supports allow supporting the second skin, by constituting, forexample, supports for depositing the compound or, when the second skinis pre-manufactured, by avoiding or limiting the creeping of the secondskin at the level of the cavities during the assembly of this secondskin.

According to a first variant, the method may comprise a step ofdepositing polymer on the TiAl-based powdery intermetallic alloy.

According to a second variant, the compound may comprise a polymer.

In each of these variants, the polymer has a binder function that allowsagglomerating the compound according to the shape that it is desired toobtain.

In order to remove the polymeric binder, the method may comprise adebinding step. The debinding may be of the thermal or aqueous type. Thedebinding step may be carried out before the consolidation step bysintering, or simultaneously.

In one form, the method may comprise a step of vertically disposing thepanel so as to inhibit creeping of the skins, in particular at the levelof the cavities of the honeycomb structure. The consolidation step, andpossibly the debinding step, can be carried out on the panel thusdisposed.

To avoid or limit the creeping of the walls of the honeycomb structurewhich delimit the cavities, walls of the honeycomb structure can be madeconvex so that, when sintering the vertically disposed panel, theseconvex walls take on a substantially planar shape under the effect ofgravity.

Of course, many other solutions can allow avoiding or limiting creepingof the walls of the honeycomb structure. As non-limiting examples, thewalls of the cavities may be corrugated and/or the panel may bevertically disposed so as to orient the cavities to increase theorientation of the walls of the cavities with respect to the horizontal.Such solutions can be combined with one another and/or with themanufacturing of a honeycomb structure with convex walls.

Further areas of applicability will become apparent from the descriptionprovided herein. It should be understood that the description andspecific examples are intended for purposes of illustration only and arenot intended to limit the scope of the present disclosure.

DRAWINGS

In order that the disclosure may be well understood, there will now bedescribed various forms thereof, given by way of example, referencebeing made to the accompanying drawings, in which:

Other features and advantages of the present disclosure will appear onreading the following non-limiting description and the appended figures,in which:

FIG. 1 is a schematic perspective view of an aircraft propulsion unitnacelle according to the teachings of the present disclosure;

FIG. 2 is a schematic perspective view of an acoustic panel of anaircraft propulsion unit nacelle according to the teachings of thepresent disclosure;

FIG. 3 is a schematic view of an acoustic panel of an aircraftpropulsion unit nacelle in a vertical position according to theteachings of the present disclosure;

FIG. 4 is a schematic perspective view of an acoustic panel of anaircraft propulsion unit nacelle in a vertical position, with wallshaving a convex honeycomb structure according to the teachings of thepresent disclosure;

FIG. 5 is a schematic front view of the panel of FIG. 4;

FIG. 6 is a partial schematic perspective view of an acoustic panel ofan aircraft propulsion unit nacelle, the panel comprising an auxeticsupport housed within a cavity according to the teachings of the presentdisclosure;

FIG. 7 is a schematic view of the auxetic support illustrated in FIG. 6;

FIG. 8 is a partial schematic perspective view of an acoustic panel ofan aircraft propulsion unit nacelle with walls having a corrugatedhoneycomb structure according to the teachings of the presentdisclosure;

FIG. 9 is a schematic perspective view of the panel of FIG. 8, the panelcomprising an auxetic support housed in a cavity according to theteachings of the present disclosure; and

FIG. 10 is a schematic view of the auxetic support illustrated in FIG.9.

Identical or similar elements are identified by identical referencenumerals in all the figures.

The drawings described herein are for illustration purposes only and arenot intended to limit the scope of the present disclosure in any way.

DETAILED DESCRIPTION

The following description is merely exemplary in nature and is notintended to limit the present disclosure, application, or uses. Itshould be understood that throughout the drawings, correspondingreference numerals indicate like or corresponding parts and features.

The teachings of the present disclosure provide a method formanufacturing a panel of an aircraft propulsion unit nacelle, forexample a panel constituting a portion of an exhaust nozzle.

An example of an aircraft turbojet engine (not represented) nacelle 1 isillustrated in FIG. 1. The nacelle 1 comprises a pylon 2 intended to befastened to a wing (not represented) of the aircraft. The nacelle 1comprises an upstream section 11 with a lip 110 forming an air inlet.The upstream section 11 is adapted to enable enhanced capture towardsthe turbojet engine of the air needed to supply a fan (not represented)and inner compressors (not represented) of the turbojet engine. Thenacelle 1 also includes a middle section 12 receiving the fan as well asa downstream section 13. Under the pylon 2 and downstream of theturbojet engine, the nacelle 1 comprises an exhaust nozzle 14 includinga gas ejection cone 141 and a primary nozzle 142. The ejection cone 141and the primary nozzle 142 of the exhaust nozzle 14 define a passage fora hot air flow exiting the turbojet engine.

The nacelle 1 and in particular the exhaust nozzle 14 may comprisepanels manufactured by a method in accordance with the teachings of thepresent disclosure, whose implementation examples are describedhereinbelow.

In general, the panels thus manufactured may constitute sub-elements ofthe nacelle 1 such as self-supporting panels and/or with or without anacoustic treatment function and/or a structural function.

In at least one form of the present disclosure the method is implementedto manufacture a panel including a honeycomb structure and skinsclasping the honeycomb structure. Such a honeycomb structure comprisescavities which may constitute, in some variations, Helmholtz resonators.

Thus, such panels may constitute all or part of a sub-element of thenacelle 1, for example all or part of the ejection cone 141, or of theprimary nozzle 142, or of an inner fixed structure 131 of the nacelle 1.These examples are in no way limiting and the teachings of the presentdisclosure may be implemented to manufacture any nacelle panel 1, forexample a panel having a complex geometry and subjected to significantthermal stresses such as parts in direct or indirect contact with thehot air flow exiting the turbojet engine.

In one form, the method implements additive manufacturing steps, beingunderstood that an additively manufactured panel may, according to theteachings of the present disclosure, be assembled with another partmanufactured by any method, for example a subtractive method.

In one form, the method comprises a step of depositing a compoundcomprising a TiAl-based powdery intermetallic alloy.

This compound may take on the form of granules or wires.

In some variations, the method does not comprise any step of melting thecompound.

In at least one variation the compound is consolidated by sintering.

In variations where a polymeric binder is used, the polymer may eitherbe deposited on the TiAl-based powdery intermetallic alloy, or it may beintegrated directly into the compound before depositing the compoundcomprising the polymer.

In some variations, the removal of the polymer may be carried out duringa specific debinding step.

The effective implementation of these different steps falls within thegeneral skills of those skilled in the art, specialized in themanufacture of nacelles by an additive method. The combination, on theone hand, of the deposition of a compound followed by consolidation bysintering - without melting - and, on the other hand, of the use of TiAlas constituent of the compound, allows manufacturing simply and at alower cost a panel, in particular with a complex geometry, reduced massand having good performance at high temperature.

FIG. 2 shows a panel 3 made using the method according to the teachingsof the present disclosure.

The panel 3 comprises a honeycomb structure 30 clasped between two skins31 and 32.

In this example, the honeycomb structure 30 constitutes Helmholtz-typecavities 34 delimited by transverse walls 33.

The skin 32 is herein an acoustic skin comprising perforations 321communicating the cells 34 with an external volume of the panel 3.

The additive manufacturing method described hereinabove may beimplemented to manufacture all or part of such a panel 3.

For example, it is possible to manufacture the skins 31 and 32 and thehoneycomb structure 30 by additive manufacturing. Alternatively, it ispossible to manufacture only the honeycomb structure 30 and one of theskins, for example the skin 31, by additive manufacturing and thenaffixed over the honeycomb structure 30 the other skin 32 previouslymanufactured by any method. Thus, in the latter case, the skin 32 may bemanufactured according to an additive method or according to asubtractive method—in the latter case, the perforations 321 may forexample be obtained by drilling a solid sheet metal before or afteraffixing this sheet metal over the honeycomb structure 30. Moregenerally, the skins and the honeycomb structure of the panel 3 may bemade according to any technique.

In one form illustrated in FIG. 3, the panel 3 may be disposedvertically to carry out the consolidation step by sintering, so as toavoid the creeping of the skins 31/32 at the level of the cavities 34 ofthe honeycomb structure 30, that is to say outside the areas where theskins are supported by the walls 33.

To avoid creeping of the cavity walls 33 during sintering with thevertically disposed panel, the honeycomb structure 30 may bemanufactured with convex walls 33 so that, when sintering the verticallydisposed panel, these convex walls take on a substantially flat shapeunder the effect of gravity.

FIGS. 4 and 5 show a panel 3 with a skin 31 and a honeycomb structurecomprising a set of horizontal and vertical walls 33 delimitingcavities, the panel 3 being represented in said vertical position. Inthis example, the horizontal walls are likely to creep during sintering.For purely illustrative reasons, these figures show two series S1 and S2of convex horizontal walls and a series S3 of planar horizontal walls.

FIG. 8 shows a panel 3 with a skin 31 and a honeycomb structurecomprising corrugated walls 33. The corrugation of the walls 33 alsoallows limiting the creeping thereof during sintering. In anon-represented form, it is possible to make walls 33 that are bothconvex and/or corrugated.

Depending on dimensions and mechanical and thermal stresses to which thepanel 3 is subjected during sintering, the dimensions and the shape ofthe convex walls can be adapted (compare for example, the series S1 withthe series S2) so that, after sintering, these walls take on asubstantially planar shape.

According to the teachings of the present disclosure, the methodcomprises a step of disposing supports 4 in respective cavities of thehoneycomb structure 30.

An example of a support 4 is shown in FIG. 7, and the disposition ofsuch a support 4 in a cavity is illustrated in FIG. 6.

Such a disposition of supports 4 within the honeycomb structure allows,for example, supporting the pre-manufactured skin 32 when it is affixedover the honeycomb structure 30, or else supporting the compound whenthe skin 32 is manufactured according to the method disclosed herein.

In order to be able to remove the supports 4 which are enclosed in thehoneycomb structure after assembling or manufacturing the skin 32, thesesupports 4 are made of a fugitive material, for example made of athermoplastic material.

Removal of the fugitive material, for example during sintering, may becarried out according to any known technique, for example by evacuationthrough the perforations 321 made in the skin 32 (see FIG. 2) or elsevia specific channels (not represented).

Such supports 4 may either be introduced into the cavities beforemanufacturing or assembling the skin 32 or be made by additivemanufacturing simultaneously with the manufacturing of the skin 31and/or of the honeycomb structure 30.

The support 4 illustrated in FIGS. 6 and 7 is auxetic so that, under theeffect of an increase in temperature, its dimension H1 between the twoskins 31 and 32 remains below a predetermined value which is selected inpractice so as to limit or prevent the expansion of the support 4 andexert a force on the skins 31 and 32 which may deform them.

In the particular case of the support 4 of FIG. 7, the support 4 has ahexagonal re-entrant shape. This support 4 comprises a lower portion 41intended to come into contact with one of the skins (skin 31), an upperportion 42 intended to come into contact with the other skin (skin 32),arms 43 and 44 having ends intended to come into contact with respectivewalls 33, and connection portions 46-48 connecting the arms 43 and 44 tothe lower 41 and upper 42 portions.

The connection portions 46-48 form, together with the lower 41 and upper42 portions, said hexagonal re-entrant shape (see FIG. 7).

Such a support 4 is dimensioned so that the expansion of the arms 43 and44 enclose the angle a formed by the connection portions 45 and 46 onthe one hand, and by the connection portions 47 and 48 on the otherhand.

It is therefore understood that the expansion of the support 4 thusallows keeping the dimension H1, which represents the dimension of thesupport 4 between the two skins, substantially constant.

FIG. 9 shows a panel 3 similar to that of FIG. 8 with corrugated walls33. In this case, it is possible to use a support 4 as illustrated inFIG. 10, comprising a corrugated periphery, and dimensioned to conformto the corrugated contours of a cavity (see FIG. 9).

Of course, the teachings of the present disclosure are not limited tothe examples that have just been described and many adjustments may bemade to these examples without departing from the scope of the presentdisclosure.

Unless otherwise expressly indicated herein, all numerical valuesindicating mechanical/thermal properties, compositional percentages,dimensions and/or tolerances, or other characteristics are to beunderstood as modified by the word “about” or “approximately” indescribing the scope of the present disclosure. This modification isdesired for various reasons including industrial practice, material,manufacturing, and assembly tolerances, and testing capability.

As used herein, the phrase at least one of A, B, and C should beconstrued to mean a logical (A OR B OR C), using a non-exclusive logicalOR, and should not be construed to mean “at least one of A, at least oneof B, and at least one of C.”

The description of the disclosure is merely exemplary in nature and,thus, variations that do not depart from the substance of the disclosureare intended to be within the scope of the disclosure. Such variationsare not to be regarded as a departure from the spirit and scope of thedisclosure.

1. A method for manufacturing a panel for a nacelle of an aircraftpropulsion unit, the panel comprising a honeycomb structure formingcavities, a first skin and a second skin clasping the honeycombstructure, the method comprising: disposing one or several supports inrespective cavities of the honeycomb structure, the supports being madeof a fugitive material such as a thermoplastic, wherein the supports areauxetic such that under the effect of an increase in temperature, adimension between the first and the second skins remains below apredetermined value.
 2. The manufacturing method according to claim 1further comprising producing the first skin and the honeycomb structureby additive manufacturing.
 3. The manufacturing method according toclaim 2, wherein the supports are made by additive manufacturingsimultaneously with at least one of the making of the first skin and thehoneycomb structure.
 4. The manufacturing method according to claim 2,wherein making the first skin and the honeycomb structure comprises adeposition of a compound comprising a TiAl-based powdery intermetallicalloy.
 5. The manufacturing method according to claim 4, wherein makingthe first skin and the honeycomb structure comprises consolidation ofthe compound by sintering.
 6. The manufacturing method according toclaim 1, 5, wherein the second skin is made by depositing a compoundcomprising a TiAl-based powdery intermetallic alloy.
 7. Themanufacturing method according to claim 1, wherein the second skin ispre-manufactured and affixed against the honeycomb structure.